In order to increase the efficiency of gas turbine power plants, both mobile and fixed, there usually must be a concomitant increase in the operating temperatures and pressures of these devices. Components made from superalloys and coated materials have allowed increased operating parameters.
By the same token, cooling air has allowed these units to operate at higher turbine inlet temperatures. Air cooling has permitted a rise in advanced turbine design inlet temperatures from 1100.degree. C. (2012.degree. F.) for uncooled blades to 1450.degree. C. (2542.degree. F.) for air cooled blades.
In some designs, the air is exhausted through many small holes in the blade, the blade root, the vane or the vane root. For the purpose of discussion, unless otherwise indicated the terms "blade" and "vane" may be used interchangeably. The cooling air, cooler than the hot expanded turbine gas, provides film cooling as well as direct internal cooling of the blade. In other designs, the cooling air is internally routed through the body of the blade. Examples of these designs may be found in U.S. Pat. Nos. 4,415,310; 3,275,294; 4,040,767; 3,909,412; 3,782,852; 3,584,458; 2,618,120; 3,647,313; and 2,487,514. Other designs are developed in Canadian patent 991,829 and U.K. patent 602,530. The aforementioned U.K. patent utilizes thermal barrier coatings and exhausts the cooling air from the trailing edge.
Current standard uncooled turbines usually operate at about 930.degree. C. (1706.degree. F.). Cooled blades, vanes (or stators) and discs operate in the 1316.degree.-1450.degree. C. (2400.degree. F.-2642.degree. F.) range. Cooling air is bled from the compressor and routed into and around the blades and vanes. Cooling is accomplished by film, transpirational and convective modes.
Current designs have a drawback in that the cooling air exits into a relatively high pressure gas stream. This requires the full compressor pressure to be used for the cooling air. Also, any exposed holes in the blade or root of the blade that has a thermal barrier coating can lead to premature failure of the ceramic coating. The degree of cooling of the blade is mainly a function of the mass flow rate of the cooling air that flows past it and is not particularly affected by the pressure of the air. It has been determined that the performance of the blades with thermal barrier coatings are limited by the cooling air. What is needed to push the gas turbine to higher performances is to use a thermal barrier coating on the blades and vanes and to change the internal air cooling system and integrate it with the turbine system.
U.S. Pat. No. 4,900,640, commonly assigned, discloses the concept of using a ceramic thermal barrier coating on a controlled expansion alloy with a coefficient of thermal expansion (CTE) such that it approximately matches the CTE of the overlaying ceramic. With the matched CTE's, the ceramic does not spall off the metal during thermal cycling. Use of the matched CTE's also allows a thicker ceramic with better insulating properties to be used than was previously the case with unmatched CTE's. The thicker thermal barrier coatings accompanied by new internal cooling arrangements disclosed and claimed here can lead to improved turbine performance.